1. Field of the Invention
The present invention relates to a portable machining system, and more particularly, to a portable fastening system for use in the manufacture of mechanical structures.
2. Background of the Invention
Traditional manufacturing techniques for assembling components to produce large mechanical structures to a specified contour have relied on fixtured tooling techniques utilizing assembly jigs and templates to locate the parts correctly relative to one another. Unfortunately, this method often yielded parts outside of acceptable tolerance because of imperfections in the templates or changes in the fixtured tooling caused by temperature variations.
To solve the problems encountered by traditional techniques, a system and method for assembling components was developed that utilized spatial relationships between key features of subassemblies as represented by coordination holes drilled into the subassemblies using numerical part definition records. The subassemblies were made intrinsically determinate of the dimensions and contour of the assembly.
The use of key features to determine the dimensions and contour of an airplane fuselage section is shown in FIG. 1. Here, a skin 20 has a plurality of stringers 22 and a plurality of shear ties 24 riveted thereon. A frame member 30 having a curved contour which is the same as the desired contour of the airplane fuselage is then riveted to the shear ties 24 and stringer clips 26.
The stringers 22, the shear ties 24 and the stringer clips 26 must be fastened to the fuselage skin 20 with extreme accuracy and consistency. Accuracy of parts manufacture ensures that the airplane will come together perfectly with no pre-stressed parts and no cosmetic imperfections.
Initially, a computer numerically controlled (CNC) machine tool performs machining operations on the skin 20. Coordination holes are drilled in the skin 20 and the stringers 22. Corresponding coordination holes are also drilled in the shear ties 24 and the stringer clips 26. A final machining operation of edge routing is performed by a high speed routing end-effector to route the edges of the fuselage skin 20 to the correct dimensions, as specified by the original part definition data base, by accurately locating the edges of the skin relative to the coordination holes in the skin.
The stringers 22 are tack fastened to the skin 20 through their aligned coordination holes. Then the shear ties 24 and stringers 22 are drilled and riveted to the skin 20. The stringer clips 26 are inserted at the correct location and are held in place while drilled and riveted to form a panel 34.
The skin 20 also has a series of panel-to-panel coordination holes 32 drilled along the edge of the skin 20. The panel-to-panel coordination holes 32 are used to position the panels 34 relative to each other. The panels 34 are still relatively flexible so the ultimate configuration is determined by the parts and their matched coordination holes. The panel-to-panel coordination holes 32 are aligned on adjacent holes and sealant is applied between the facing surfaces of the panel edges. The panels 34 are aligned so that the panel-to-panel coordination holes 32 on adjacent panels 34 line up exactly and the two panels are fastened together at their adjacent edges by temporary deco fasteners through the coordination holes. The panels are then drilled and riveted to permanently fasten them together to form a super panel 36.
Coordination holes are drilled into the frames 30 and are aligned with the coordination holes in the stringer clips 26. The frames 30 are fastened and their alignment determines the contour of the super panel 36. Thus, the contour is independent of any hard tooling. Once the super panel 36 is formed, the temporary cleco fasteners holding the parts in position are replaced by permanent fasteners.
The super panels 36 are temporarily fastened using the panel-to-panel coordination holes 32 to form fuselage quarter panels which are in turn temporarily fastened to form a lower fuselage lobe 38A and an upper fuselage lobe 38B, as shown in FIGS. 2A and 2B. A floor grid 40 is aligned with the lower lobe 38A using coordination holes, and is fastened in place. The fixture 44 does not determine the contour or dimensions of the fuselage. Instead, the coordination holes drilled into the floor grid 40 determines the cross-dimensions of the fuselage 42.
Once the frame members 30 and lobe skin coordination holes 46 are all aligned and temporarily fastened with deco fasteners, they are drilled to form the final fuselage section 42, as shown in FIG. 2B. The fuselage section 42 is then disassembled, deburred, cleaned, and sealant is added.
After sealing, each super panel 36 is again aligned using the coordination holes. The overlapping portion of the panels 36, a lap joint 48, is shown in FIGS. 2B and 2C. Each lap joint 48 has a plurality of columns 50, where each of the columns 50 has 3 rows of rivets 52A-C. Two rivets of the rows 52A and 52C are for rivets that require a countersink as well as drilling.
The super panels 36 could be fastened to form a quarter panel by an assembly device, such as that described in U.S. Pat. No. 4,662,556 (the '556 patent). However, the device described in the '556 patent moves a working unit along a guide beam that is supported by two huge arc-shaped girders, and could not be used to form the lower or upper fuselage lobes 38A and 38B, respectively, because of its size and design. Simply put, the unit described in the '556 patent or any variations thereof would not fit within the fuselage lobes 35A and 38B, and certainly not the fuselage assembly 42. Attempts to redesign the assembly device discussed in the '556 patent to handle larger portions of the fuselage assembly 42 have failed because of severe problems with vibration which interfered with the proper seating of fasteners such as rivets. Further, the assembly device discussed in the '556 patent is not versatile and requires an expensive and heavy support structure.
Presently, the fuselage quarter panels 36 and, lower and upper lobes 38A and 38B, and the final fuselage assembly 42 are re-tacked into position after being filed, cleaned, and sealed. Then, the panels 36 are riveted together by hand, where one person stands on a platform (not shown) outside the fuselage, inserting and then pneumatically driving a rivet fastener while another person stands inside the fuselage, bracing a large bucking bar against a rivet shank and holding it in place by leaning against the bucking bar with his shoulder. Obviously, such a process presents a risk of injury. Further, the manual process results in rivets that were unevenly deformed, poorly seated, or riveted too close to an edge of the lap joint 48.
Unfortunately, the manual process is dangerous, time-consuming, expensive and often leads to extensive rework. Consequently, there is a need in the art for a fastening system that speeds up production, ensures riveting and drilling accuracy, eliminates the required step of disassembling the entire fuselage to de-burr, clean and seal, and can be operated within the final fuselage assembly 42.